Rocketdyne J-2

The J-2, commonly known as Rocketdyne J-2, was a liquid-fuel cryogenic rocket engine used on NASA's Saturn IB and Saturn V launch vehicles. Built in the United States by Rocketdyne, the J-2 burned cryogenic liquid hydrogen (LH₂) and liquid oxygen (LOX) propellants, with each engine producing 1,033.1 kN (232,250 lb_f) of thrust in vacuum. The engine's preliminary design dates back to recommendations of the 1959 Silverstein Committee. Rocketdyne won approval to develop the J-2 in June 1960 and the first flight, AS-201, occurred on 26 February 1966. The J-2 underwent several minor upgrades over its operational history to improve the engine's performance, with two major upgrade programs, the de Laval nozzle-type J-2S and aerospike-type J-2T, which were cancelled after the conclusion of the Apollo program.

The engine produced a specific impulse (I_sp) of 421 seconds (4.13 km/s) in a vacuum (or 200 seconds (2.0 km/s) at sea level) and had a mass of approximately 1,788 kilograms (3,942 lb). Five J-2 engines were used on the Saturn V's S-II second stage, and one J-2 was used on the S-IVB upper stage used on both the Saturn IB and Saturn V. Proposals also existed to use various numbers of J-2 engines in the upper stages of an even larger rocket, the planned Nova. The J-2 was America's largest production LH2-fuelled rocket engine before the RS-25. A modernized version of the engine, the J-2X, was considered for use on the Earth Departure Stage of NASA's Space Shuttle replacement, the Space Launch System.

Unlike most liquid-fueled rocket engines in service at the time, the J-2 was designed to be restarted once after shutdown when flown on the Saturn V S-IVB third stage. The first burn, lasting about two minutes, placed the Apollo spacecraft into a low Earth parking orbit. After the crew verified that the spacecraft was operating nominally, the J-2 was re-ignited for translunar injection, a 6.5 minute burn which accelerated the vehicle to a course for the Moon.

The J-2's thrust chamber assembly served as a mount for all engine components, and was composed of the thrust chamber body, injector and dome assembly, gimbal bearing assembly, and augmented spark igniter.

The thrust chamber was constructed of 0.30 millimetres (0.012 in) thick stainless steel tubes, stacked longitudinally and furnace-brazed to form a single unit. The chamber was bell-shaped with a 27.5:1 expansion area ratio for efficient operation at altitude, and was regeneratively cooled by the fuel. Fuel entered from a manifold, located midway between the thrust chamber throat and the exit, at a pressure of more than 6,900 kPa (1,000 psi). In cooling the chamber, the fuel made a one-half pass downward through 180 tubes and was returned in a full pass up to the thrust chamber injector through 360 tubes. Once propellants passed through the injector, they were ignited by the augmented spark igniter and burned to impart a high velocity to the expelled combustion gases to produce thrust.

The thrust chamber injector received the propellants under pressure from the turbopumps, then mixed them in a manner that produced the most efficient combustion. 614 hollow oxidizer posts were machined to form an integral part of the injector, with fuel nozzles (each swaged to the face of the injector) threaded through and installed over the oxidizer posts in concentric rings. The injector face was porous, being formed from layers of stainless steel wire mesh, and was welded at its periphery to the injector body. The injector received LOX through the dome manifold and injected it through the oxidizer posts into the combustion area of the thrust chamber, while fuel was received from the upper fuel manifold in the thrust chamber and injected through the fuel orifices which were concentric with the oxidizer orifices. The propellants were injected uniformly to ensure satisfactory combustion. The injector and oxidizer dome assembly was located at the top of the thrust chamber. The dome provided a manifold for the distribution of the LOX to the injector and served as a mount for the gimbal bearing and the augmented spark igniter.

The augmented spark igniter (ASI) was mounted to the injector face and provided the flame to ignite the propellants in the combustion chamber. When engine start was initiated, the spark exciters energized two spark plugs mounted in the side of the combustion chamber. Simultaneously, the control system started the initial flow of oxidizer and fuel to the spark igniter. As the oxidizer and fuel entered the combustion chamber of the ASI, they mixed and were ignited, with proper ignition being monitored by an ignition monitor mounted in the ASI. The ASI operated continuously during entire engine firing, was uncooled, and was capable of multiple reignitions under all environmental conditions.

Thrust was transmitted through the gimbal (mounted to the injector and oxidizer dome assembly and the vehicle's thrust structure), which consisted of a compact, highly loaded (140,000 kPa) universal joint consisting of a spherical, socket-type bearing. This was covered with a Teflon/fiberglass coating that provided a dry, low-friction bearing surface. The gimbal included a lateral adjustment device for aligning the combustion chamber with the vehicle, so that, in addition to transmitting the thrust from the injector assembly to the vehicle thrust structure, the gimbal also provided a pivot bearing for deflection of the thrust vector, thus providing flight attitude control of the vehicle.